АС 27-1A 7/30/97 As can be seen from the foregoing, there is no such thing as an exact allowable wind for a particular test or rotorcraft. The flight test team must decide on the allowable wind for each condition based on all available information and their engineering judgment. The following summary of allowable wind conditions is given for general guidance only:
(A) Hover performance - 0 to 3 knots.
(B) Height-velocity - 0 to 3 knots.
(C) IGE controllability and maneuverability - 0 to 5 knots.
(iv) A means should be provided to measure the wind velocity, direction, and ambient air temperature at the rotor height for any particular tests.
(2) Altitude Effects. Using FAA/AUTHORITY-approved methodology, hover performance may be extrapolated and/or interpolated from test data up to a maximum of +4,000 feet. Experience has shown that IGE handling qualities, height-velocity, and engine operating characteristics should not be extrapolated higher than approximately 2,000 feet density altitude from the test altitude. Cruise stability/controllability tests should be evaluated at least at two different altitudes, the lowest practical altitude and approximately the highest cruise altitude requested for approval. This can allow an interpolation of approximately 10,000 feet. As in all testing, extrapolation and/or interpolation should only be considered if all available information and engineering judgment indicate that regulatory compliance can be met at the untested conditions.
(3) Altitude Limitations.
Explanation.
(A) Two altitudes are normally presented in the RFM to define the operating envelope of a rotorcraft;
- Maximum operating altitude, and
- Maximum takeoff and landing altitude.
(B) Maximum operating altitude is an operating limitation required by § 27.1527 and delineates the maximum altitude to which operation is allowed. This altitude normally constitutes the maximum cruise or en route altitude.
(C) Maximum takeoff and landing altitude is the hover in-ground-effect (IGE) ceiling for a rotorcraft as described in § 27.73. The hover ceiling and any information pertinent to takeoff and landing are presented in the performance information section of the RFM. For rotorcraft certified to CAR 6, Amendment 6-7 or any amendment of FAR 27, a hover ceiling may not be presented above the altitude at which H-V and IGE controllability tests were conducted plus allowable extrapolation, Par 58 62
unless that extrapolated altitude is at least 7,000 feet. If the applicant elects to demonstrate these tests to an altitude below 7,000 feet, then that altitude is the maximum takeoff and landing altitude of the rotorcraft. The maximum takeoff and landing altitude may be coincident with, but never above the maximum operating altitude limitation. Takeoff and landing and hover ceiling data and presentation requirements are presented in §$ 27.51, 27.73 and 27.1587.
(її) Procedures.
(A) In establishing the maximum takeoff and landing altitude, the following tests are normally required:
(1) Takeoff (§27.51)
(2) Climb (§§ 27.65 and 27.67)
(3) Performance at minimum operating speed (§ 27.73)
(4) Landing (§ 27.75)
(2) Limiting height-speed envelope (§ 27.79)
(6) GE controllability (§ 27.143c) (Z) Cooling (§§ 27.1041, 27.1043 and 27.1045)
(8) Engine operating characteristics (§ 27.939) Specific guidance on test methodology and data requirements is provided in applicable paragraphs of this AC.
(B) As detailed in subparagraph b(2) above, the maximum allowable extrapolation of H-V, IGE controllability and engine operating characteristics is +2,000 feet. Therefore, the maximum takeoff and landing altitude presented in the RFM is not normally more than 2,000 feet above the density altitude experienced at the high altitude test site, or for CAR 6, Amendment 6-7 and subsequent, unless test results were demonstrated to at least 7,000 feet.
(C) If IGE controllability is demonstrated to at least 17 knots of wind at 7,000 feet, hover capability above 7,000 feet may be presented provided that the maximum demonstrated safe wind for takeoff and landing above 7,000 feet is specified in the RFM.
Par 58 63
АС 27-1А 7/30/97 (D) The requirements for data collection and presentation in the RFM vary depending upon the certification basis of the rotorcraft. These requirements are presented by regulation and amendment in Figures 58-1 and 58-2.
(E) The maximum takeoff and landing altitude may be extrapolated no greater than the values given in Paragraph b(2) and not above the lowest limiting altitude resulting from the requirements of subparagraph A of this paragraph.
(4) Temperature Effects.
Background.
(A) In the past, approved analyses were frequently accepted for determining the extreme temperature effects on performance and flight characteristics.
With the introduction of newer, higher performance rotorcraft, advanced rotor blade designs, higher airspeeds, and higher blade tip Mach numbers, the previous methods have proven to be insufficient. Therefore, the performance and flight characteristics should be validated at extreme temperatures; however, analysis may be permitted if a suitable methodology is demonstrated.
(B) Various FAA/AUTHORITY cold weather programs have verified that rotorcraft can be affected by cold temperature in both the performance and flying qualities areas. Hot temperature conditions, although not shown to be as critical for flying qualities, should be given consideration.
(C) Additionally, design deficiencies surfaced when the rotorcraft were exposed to temperature extremes and some of these difficulties were severe enough to require the redesign of equipment and/or materials. Therefore, to satisfy § 27.1309(a), the applicant needs to substantiate the total rotorcraft throughout the foreseeable range of operating temperatures.
(її) Procedures.
(A) The FAA/AUTHORITY is responsible for verifying the effects of temperature on performance and handling characteristics. A limited flight verification, if necessary, could include spot checks of hover performance, IGE controllability, vibration, simulated power failure, static stability, height-velocity, VNE/No evaluations, ground resonance, etc. In addition, systems should be evaluated to determine satisfactory operations.
(B) Extrapolation of test data should only be allowed if the applicant's predicted or calculated data is verified by actual test, but in any case extreme caution should be used for extrapolations that are 10° C below or 20° C above those values tested.
Par 58 64
7/30/97/ АС 27-1A
(5) Engine Power - Turboshaft Engine.
(i) Background.
(A) The purpose of rotorcraft performance flight testing is to obtain accurate quantitative flight test performance data to provide flight manual information.
(B) Flight tests are designed to investigate the overall performance capabilities of the rotorcraft throughout its operating envelope. This testing furnishes information to be included in the flight manual and provides a means of validating the predicted performance of the rotorcraft with a minimum installed specification engine.
(C) The power used to complete the flight manual performance must be based on power values no greater than that available from the minimum uninstalled specification engine after it is corrected for installation losses. A minimum uninstalled specification engine is one that, on a test stand under conditions specified by the engine manufacturer, will produce the certificated power at specification temperatures and/or speeds. The specification values may be either a rating or limit. Some engine manufacturers certify an engine to a specitied power at a particular engine temperature or speed rating with higher allowable limits. The limit is the maximum value the installed engine is allowed in order to develop the specification power. Prior to installation of each engine in a rotorcraft, the performance is measured by the engine manufacturer. This is done by making a static test run in a test cell and referring the results to standard day, sea level conditions. The performance parameters obtained are presented as uninstalled engine characteristics on a test log sheet. This is commonly referred to as a "final run sheet." Figure 58-3 compares a typical engine to one the manufacturer has certified as a minimum uninstalled certified engine.
(D) After engine certification, the engine manufacturer is responsible to ascertain that each engine delivered will produce, as a minimum, the certitied power without exceeding specification operating values; therefore, a "final run sheet" is created for every engine produced. Additionally, if needed, arrangements can usually be made with the engine manufacturer to obtain a torque system calibration for individual engines. This will further optimize the accuracy of the engines used in the flight test program. The engine manufacturer will also provide predicted uninstalled power available for the various power ratings. This information may be derived from an engine computer "card deck" and from charts and tables in the engine detail installation manual. These data also provide engine performance for the range of altitudes and temperatures approved for the engine and include methods for correcting this performance for installation effects. The parameters contained in a typical "card deck" are plotted for one engine rating in Figure 58-4.
(E) Several power losses may be associated with installing an engine in a rotorcraft. Typical losses are air inlet losses, gear losses, air exhaust losses, and powered accessory losses such as electrical generators. Additional flight Par 58 65
manual performance considerations are the torque indicating system accuracy and torque needle split. The predicted uninstalled power available engine characteristics cannot be assumed to be the actual power available after the engine is installed in the rotorcraft because this procedure would neglect the installation power losses. It is necessary to know the installation losses in order to determine the flight manual performance. Installation losses are reflected reductions in available power resulting from being installed in a rotorcraft. These losses usually consist of those incurred due to engine inlet and/or exhaust design. The rotorcraft manufacturer conducts tests to confirm the installed specification engine power available on which published performance is based. The specific methods used vary widely between manufacturers but usually include some combination of ground and flight tests. Figure 58-5 is a typical installed power available chart for one set of conditions.
(F) The installed power available is, in most cases, lower than obtained on a test stand. This is especially true at lower airspeeds where exhaust reingestion may occur and there are changes in airflow routing. The rotorcraft manufacturer may elect to determine the installation losses for different flight conditions to take any airspeed advantages. This is acceptable if, for example, the hover performance is based on the actual power available from an installed minimum specification engine in a hover. Likewise, it is permissible for the rotorcraft manufacturer to determine his climb performance based on the actual power available from an installed minimum specification engine at the published climb airspeed. This will allow the manufacturer to take advantage of, for example, increased inlet efficiency.
(її) Procedures.
(A) The installed minimum specification engine power output has been predicted and calculated for various flight conditions. It is imperative that the predicted values be verified by actual flight test. The flight test involves obtaining engine performance measurements at various power settings, altitudes, and ambient temperatures. The data should be obtained at the actual flight condition for which the performance is to be presented (i.e., hover, climb, or cruise).
(B) Following a power increase, engine temperature and/or RPM can significantly decrease for a period of time as torque is held constant. Said another way, torque will increase if RPM and/or temperature are held constant. This is a characteristic typical of turbine engines due largely to expansion of turbine blades and reduced clearances in the engine. Some engines may show a temperature increase at constant power due to engine or temperature sensing system peculiarities. An engine will usually establish a stabilized relationship of power parameters in approximately 2 or 3 minutes. For this reason, the following procedure should be used when obtaining in-flight engine data.
(1) To determine the takeoff and 2 ½-minute values, first stabilize the engine at a low power setting. After stabilization, rapidly increase the power Par 58 66
demand to takeoff and/or 2 ½-minute power levels. Record the engine parameters as soon as the specification torque, temperature, or speed is attained. Care must be taken not to exceed a limit. These readings should be obtained approximately 15 seconds after power is initially applied.
(2) To determine the 30-minute and/or maximum continuous power values, approximately 2 to 3 minutes of stabilization time after power is increased is generally used, but up to 5 minutes stabilization time is allowed. The reason for the different procedures is when a pilot requires takeoff or 2 ½-minute power values he is in a critical flight condition and does not have the luxury of waiting for the engine(s) to produce rated power. Stabilization time is allowed for the maximum continuous and 30-minute ratings because these values are not associated with flight conditions for which power is needed immediately.
(C) The in-flight measurements recorded with the engine(s) on the flight test rotorcraft must be corrected downward if the test engine is above minimum specification and corrected upward for a test engine that is below minimum specification. This correction is necessary to verify that a minimum specification engine installed on a production rotorcraft is capable of producing the power values used to compute the flight manual performance without exceeding any engine limit. In addition, if the production rotorcraft's power measurement devices have significant (greater than 3 percent) power error, this error must be accounted for in a conservative manner.
(D) On multiengine rotorcraft, the engine location may result in different installation losses between engines. If this condition exists, multiengine performance should be based on the total power available after considering the different installation losses and with minimum specification engines installed.
One-engine-inoperative performance must be based on the loss of the engine which has the lowest installation losses. Additionally, the power losses due to such items as accessory bleed air, particle separators, engine driven accessories, etc., must be accounted for accordingly.
(E) Power available data should be obtained throughout the test program at various ambient conditions. Some engines have devices which restrict the mechanical No speed to a constant corrected speed at cold temperatures. Others may limit power to a fuel flow value which would be encountered only at certain ambients.
Others may limit by torque limiting devices. Therefore, power available data should be obtained at various ambients to verify that all limiting devices are functioning properly and have not been affected by the installation.
(F) Through use, turbine engine power capabilities decrease with time. This is called engine deterioration. Deterioration is largely a function of the particular engine design, the manner, and the environment in which the engine is operated. There is a need, therefore, to provide a method which can be used in service to periodically determine the level of engine deterioration. A power assurance curve is Par 58 67
AC 27-1A 7130/97 usually provided to allow the flightcrew to know the power producing capabilities of any engine. A power assurance check is a check of the engine(s) which will determine that the engine(s) can produce the power required to achieve flight manual performance.
This check does not have to be done at maximum engine power. Figure 58-6 is a typical power assurance curve for an installed engine showing minimum acceptable torque which assures that power is available to meet the rotorcraft flight manual performance. Some power assurance curves have maximum allowable No limits that must not be exceeded for a given torque value. An in-flight power assurance check may be used in addition to the pretakeoff check. The validation of either check must be done by the methodology used to determine the installed minimum specification engine power available. For the in-flight power assurance check there must be full accountability for increased efficiency due to such items as inlet ram recovery, absence of exhaust reingestion, etc. A power assurance check done statically and one conducted in-flight must yield the same torque margin(s). An engine may pass power assurance at low power but still may not be capable of producing the rated power values. This occurs when the curve of corrected power and corrected temperature for the engine intersects the minimum uninstalled specification engine curve. If this condition exists, the entire power assurance and power available information may need to be reestablished.
(6) Deteriorated Engine Power - Turboshaft Engine.
Background.
(A) A specific engine model may have been certificated for operation with power which has "normally" deteriorated below specification. This "normal" deterioration refers to a gradual loss in engine performance, possibly caused by compressor erosion, as opposed to a sudden performance loss which may be due to mechanical damage. The application for deteriorated engine power should not be confused with the installed mechanical engine derating which is frequently used to match transmission and engine power capabilities.
(B) The use of deteriorated power is intended to allow continued operations with an engine which is serviceable and structurally sound, although aircraft performance may be depreciated. The useful life of the engine may, therefore, be extended at a dollar savings to the operator.
(C) Although installed performance is the primary topic in this discussion, considerations must be given to other operational characteristics and systems which may be affected by deteriorated engine power. These include:
(1) Engine characteristics (§ 27.939). Surge margin, engine response, and air-restart capability might be affected and should be addressed, but flight testing may not be required depending on the individual engine/aircraft installation and fuel scheduling mechanism.
Par 58 68
(2) Performance of customer bleed air systems may be degraded slightly. No problem would be anticipated unless certain items within the system depend on a critical Pc for their function.
(3) The maximum attainable gas producer speed, and thus power available under certain ambients, may be affected if Pc is an input to the fuel scheduling mechanism.
(4) Systems for surge protection which schedule on Pc such as bleed valves, flow fences, bleed bands, and variable inlet guide vanes may be influenced. The effect would normally be negligible unless when installed, the installation losses, combined with reduced Pc because of deterioration, would cause the bleed device to open and reduce power at any one of the engine ratings.
Procedures.
(A) The need for flight tests to verify predicted power available with deteriorated engines depends on the scope of testing which occurred during initial certification. If the original rotorcraft certification included flight testing as described in Paragraph (5) (engine power-turboshaft engines) herein for validation of power available, the need for a demonstration with deteriorated engines is greatly diminished and perhaps eliminated.
(B) If flight testing to verify deteriorated engine power available is deemed necessary, the procedure used would be the same as that described in Paragraph (5) (engine power-turboshaft engines), except that the data would be corrected downward to a deteriorated engine runline. Efforts should concentrate on obtaining data in areas of the operational envelope where maximum gas producer speed is likely to be attained, or where bleed valves or other devices which schedule on gas producer discharge pressure are likely to function. On many installations maximum gas producer speed will occur with cold temperatures and high altitudes; bleed valves and other devices which schedule on gas producer discharge pressure are most likely to function and reduce power on a hot day at low altitude.
(C) The adjustments to the normal power assurance check procedures for deteriorated engines will be influenced by the preferences of the aircraft manufacturer and by any special stipulations of the engine certification established as a condition for the engine to remain in service when below specification. Possibly, more stringent and more complicated engine monitoring procedures will be introduced when allowing the use of deteriorated power; for example, an in-flight trend monitoring program with the associated bookkeeping duties may be required. Such an in-flight procedure must be evaluated by flight tests as described in Paragraph (5) (engine power-turboshaft engines) herein. Normally, however, the manufacturer would be expected to present a modification, or extension of the power assurance procedure Par 58 69
already in place for the specification engine, which could eliminate the need for flight test evaluation.
58A. § 27.45 (Amendment 27-21) GENERAL.
a. Explanation. Amendment 27-21 adds § 27.45(f) to the regulation. This section establishes the requirement for furnishing power assurance information for turbine powered rotorcraft. This information is to provide the pilot a means of determining, prior to takeoff, that each engine will produce the power necessary to achieve the performance presented in the rotorcraft flight manual (RFM).
b. Procedures. All of the policy material pertaining to this section remains in effect. In addition, the power assurance information included in the RFM should be verified. Although this requirement is normally met with a power assurance curve, other methods of compliance may be proposed.
58B. § 27.45 (Amendment 27-29) GENERAL.
a. Explanation. Amendment 27-29 added the requirements for certification of 30-second/2-minute One Engine Inoperative (OEI) power ratings. For rotorcraft approved for the use of 30-second/2-minute OEI, partial power checks currently accomplished with approved power assurance procedures for lower power levels may not be sufficient to guarantee the ability to achieve the 30-second power level.
b. Procedures. Information provided in Appendix 1 of this AC includes guidance material on power assurance procedures to assure that the OEI power level can be achieved.
Par 58 70
CERTIFICATION BASIS Rgmts 27-Amdt. 21 27-Amdt. 2 FAR 27 Original 6-Amdt. 7 6-Amdt. 4 CAR 6 Original Original CAR 06 7/30 / 97
1. MGW Sea 1. MGW Sea 1. MGW Sea 1. MGW Sea 1. MGW Sea 1. MGW Sea 1. No specific H-V Level Level Level Level Level Level wt. and alt.
00 Ref. 2. Max. OGE 2. Wt. selected 2. Wt. selected 2. Wt. selected 2. Wt. selected 2. Wt. selected requirements.
27.25 wt. Lesser of: by applicant at by applicant at by applicant by applicant by applicant at
27.1519 27.79 Test b. 7000' Hd a. Max. alt. cap. lessor of: a. Max. alt. cap. lessor of: a. Max. alt. cap. a. Max. alt. cap. at lessor of: Idg. alt. max. t.o. and Idg. alt. max. t.o. and
27.1587 Conditions b. 7000' Hd b. 7000' Hd b. 7000' Hd
6.116
6.741
6.743
3. H-V is perf. 3. H-V is perf. 3. H-V is perf. 3. H-V is perf. 3. H-V is perf. 3. H-V is H-V is Info. info. info. info. info. operating operating
4. Max. alt. For 4. If H-V wt is limitation. limitation.
REM which H-V is valid less than IGE Wt., wt. becomes limitation for
t.o.Adg.
AC 27-1 5. If H-V is less Para than OGE wt., 58 & 59 H-V wt. becomes limit.
6. Applicant is encouraged to demo H-V to Remarks WAT limits.
7. Hover data may be shown above 7000' if AC 27-1A H-V & IGE are demo'd to 7000' FIGURE 58-1. H-V REQUIREMENTS
CERTIFICATION BASIS FAR 27 CAR 06 Ramts 27-Amdt. 21 Original Original Original
1. MGW Sea Level 1. MGW Sea Level 1. MGW Sea Level 1. No specific CONTROL 27.1587 27.143 27.25 IGE Ref. Conditions Test b. 7000' Hd 2. Max. IGE wt. Lesser of: a. Max. alt. cap. 3. Critical CG 4. Critical Rotor t.o. and Idg. alt. 2. Wt. selected by 3. Critical CG applicant to max. 4. Critical Rotor RPM 2. Max. approved wt. for t.o./dg at alt 3. Critical CG 4. Critical Rotor RPM 5. Wind not less than above sea level. * get ayed in requirement. requirement.
AC 27-1 RPM 5. Critical wt. 20 mph.
Para 5. Wind of not less 6. Wind of not less 58 & 81 than 17 kts. than 20 mph REM 6. Max. safe wind is 7. Max. safe wind 6. Max. safe wind is perf. info. is perf. info. perf. info.
7. If 17 kts. wind demo'd to alt. less Remarks than 7000', a corresponding WAT limit must be established.
7/30/97 FIGURE 58-2. IGE CONTROLLABILITY REQUIREMENTS
600 CORRECTED SHAFT HORSEPOWER 500 ENG S/N 12345- 400 300 MIN CERTIFIED
SPECIFICATION ENGINE 200 100 1000 1100 1200 1300 1400 1500 CORRECTED TOT °F FIGURE 58-3. SHAFT HORSEPOWER VS TURBINE OUTLET TEMPERATURE -
SEA LEVEL STANDARD DAY Par 58 73